1. Field of the Invention
The present invention relates to the field of turbomachines and, in particular, to that of gas turbine engines, such as turbojet engines, turboprop engines or engines with high speed fans (also known as “propfans”). It is aimed at a vibration damping device for the blade attachments of these engines.
2. Description of the Related Art
Aeronautical turbomachines are made up of a plurality of bladed rotors, that is to say of rotary disks to the peripheries of which moving blades are attached. These bladed rotors are particularly sensitive components because in terms of their design they have to meet requirements of mechanical integrity when rotating and when under aerodynamic load. All of these aspects mean that these structures are statically loaded and, given the life requirements, the amplitudes of vibrations that they experience need to remain small.
The design and development of a turbomachine involves coordinating several disciplines which means that the design process is an iterative one. Vibrational design is carried out in order to avoid the presence of critical modes in the operating range. The whole is validated at the end of the design cycle through an engine test in which vibrational amplitudes are measured. Sometimes high induced levels arise either as a result of synchronous or asynchronous forced responses or as a result of instabilities. The design has then to be reviewed, this being a process which is particularly lengthy and expensive.
The objective from an industrial standpoint is therefore to predict, as early on as possible in the design cycle, what the levels of vibrational response of the structures will be so that the required corrective measures can be taken as early on as possible in the design process. Mechanical damping, which is included in this category, is an important aspect for the designers to address.
The damping of compressor airfoils is a special problem that needs careful attention because these airfoils are particularly sensitive to vibrational phenomena, especially when their length is great. This problem is therefore particularly acute in respect of the airfoils of the first stage of the low-pressure compressor, whether this is a turboprop stage, the bladed rotor of which is not ducted, a bypass turbojet stage, the rotor or “fan” of which is ducted, or even an unducted rotor of a propfan engine.
It is also particularly tricky in the case of propfan engines because, on the one hand, these airfoils are twice as slender as the current airfoils used in a fan and therefore more sensitive to phenomena of flutter and, also, the fact that there are two rows of contrarotating fans produces significant forced excitation stresses on account of the wake effect that the first fan has on the second. Coupling between the vibration modes of the two rows of contrarotating fans through the structure that supports them and which may prove destructive to the engine are also sometimes encountered. In addition, propfans, unlike ducted fans, are sensitive to loadings known as 1P loadings which arise when the engine adopts an angle of incidence, notably when the airplane turns on takeoff. During these phases, the airfoil of a propfan does not experience an even angle of incidence of the air stream as this varies according to its angular position and is therefore subjected to specific excitation synchronous with the speed of the engine.
The airfoils are conventionally attached to the compressor disk by assemblies of the pinned attachment type, that is to say by open cavities into which bulbs that form the blade roots are slid. These cavities are cut into the disk and have retaining walls against which the corresponding faces of the blade root bear.
Devices for reducing blade vibrations have been designed, one example being the one described in NASA U.S. Pat. No. 6,102,664, and which involves bonding a viscoelastic material onto those faces of the blade root that are in contact with the retaining walls of the cavities in the disk. This technique has the disadvantage of requiring a modification to the method of manufacture of the fan or propfan blades and of not being suited to retrofitting to existing blades. It also has the disadvantage that the entire blade has to be changed if there is deterioration of the damping device, unlike a configuration in which the damping device is separate from the blade as proposed here.
Another technique from the prior art involves inserting a shim between the surface of the cavity and that of the blade root bulb. Such a shim, described in General Electric Company U.S. Pat. No. 5,240,375, takes the form of several metal layers assembled as a sandwich, with a layer of austenitic steel sandwiched between two layers of phosphor bronze that have a low coefficient of friction. However, it is aimed at avoiding wear of contacting components and has no appreciable impact on the ability of the airfoils to withstand vibrational stresses.
The applicant company's patent EP 2014873 describes a shim with rigid layers alternating with layers made of an elastic material.